There is a continual need to decrease the fuel consumption of aircraft gas turbine engines (for example, in terms of Specific Fuel Consumption (SFC)), in order to save operating costs, and to reduce their environmental impact due to carbon emissions and nitrous oxide (NOx). Another requirement is to reduce “contrails” formed by condensing water vapour produced by aircraft engines when operating at high altitude in some circumstances, again in order to reduce the impact of the aircraft on the environment.
Gas turbine engines comprising “sequential” combustors have been proposed for use in electricity production. Examples include the GT24 gas turbine engine produced by Alstom™. In such gas turbine engines, a second combustor is provided downstream of a first combustor between upstream and downstream turbines. Such an arrangement is described for example in U.S. Pat. No. 5,941,060. Such a system comprises a pair of “constant pressure” combustors, in which fuel is burnt at approximately constant pressure between turbine stages. This fuel is then expanded in a separate downstream turbine stage.
Sequential combustors have also been proposed for aircraft gas turbine engines, particularly for military aircraft, for example in “Two-combustor Engines' Performances under Design and Off-design Conditions” by A S Lee, R Singh and S D Probert, presented at the 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit in August 2009. However, previous studies have only found arrangements which can be expected to have increased SFC compared to gas turbines having conventional combustors.
This problem is particularly pronounced in gas turbine engines which are required to operate efficiently across a wide range of thrust/power levels. Generally, gas turbine engines have a “design point” at or close to their maximum power at which they produce power most efficiently. At power levels below this design point (so called “off design conditions”), power is produced significantly less efficiently. This is the case for example in gas turbine engines used in civil aircraft, in which the required thrust varies across different phases of flight, and also in gas turbine electrical generators or gas turbines in use in ships, in which reduced power may be required in some circumstances. Single spool gas turbine engines can operate at reduced power relatively efficiently by varying the position of inlet guide vanes upstream of the compressor, thereby allowing operation at a substantially constant turbine entry temperature across a wide range of conditions. However, such operation is not generally practical in multi-spool gas turbines having several compressors driven by separate turbines, as this would require variable geometry turbines or exhaust.
There is also increasing concern in the field of aviation regarding vapour trails produced by high flying aircraft. Vapour trails are artificial clouds that are visible trails of condensed water vapour exhausted by vehicles' engines. Vapour trails may be formed as warm, moist exhaust gas mixes with ambient air, and arise from the precipitation of microscopic water droplets or, if the air is cold enough, ice crystals. It is known that, depending on the timescale considered, the climate-warming impact of aircraft exhaust vapour trails and resulting vapour trail cirrus is of a magnitude similar to or perhaps even greater than that of the CO2 emitted by aircraft, and therefore represents a significant element of aviation's total climate impact. It is therefore desirable to reduce or eliminate the formation of vapour trails from aircraft.
Engines in which further fuel is added between turbine stages are also known (inter-turbine burning), as described for example in GB 514620. In such a cycle, partial combustion takes place in the combustor, and the mixed combustion products and unburned hydrocarbons are passed to a turbine. In the turbine, further fuel is added and burnt and simultaneously expanded (inter-turbine burning), such that the expansion through the turbines is substantially at a constant temperature. Such a process is known as an “isothermal” process, in contrast to the constant pressure process described above, where the fuel is burn completely as feasible in the combustors. Consequently, in a constant temperature process, the expansion within the turbine is closer to an isothermal process while in a gas turbine having sequential combustors, the expansion in the turbines is adiabatic. However, such isothermal expansions are inefficient and such cycles require the use of a recuperator (i.e. a device to transfer heat from the exhaust to the compressor outlet) to overcome the inefficiency of such isothermal cycles.
The present invention describes a gas turbine engine and a method of operating a gas turbine engine which seeks to overcome some or all of the above problems.